This invention relates to composite structures, and, more particularly, to the fabrication of a structure having a laminated skin structure bonded to a substrate.
Multilayer laminated composite structures are used in a variety of applications requiring high strength and low weight. For low and moderate temperature applications, the structures are typically made of light-weight composites of fibers embedded in an organic matrix. In some cases, substrates such as structural foams are bonded into the structures as well. Through careful selection of materials and processing, this composite-design approach offers the opportunity to optimize many of the properties of the structure.
Such structures are generally fabricated from a number of individual elements carefully selected to achieve the required performance objectives, and thereafter bonded together. Some or all of the components may initially be in an uncured state, so that curing and post-curing steps are used in the fabrication procedure. The curing and post-curing steps are performed by heating the structure according to a temperature-time schedule specified to cure and, optionally, post-cure the organic components of the composite material.
Although the final fabricated structure may have exceptional performance, the fabrication operation may present challenging problems. One of the ongoing obstacles to the fabrication of laminated composite structures is a consequence of the differing coefficients of thermal expansion of the constituents of the composite material. For example, if two components having differing coefficients of thermal expansion are bonded together and then heated to elevated temperature for curing, thermal strains and stresses are created within the cured structure upon cooling. When there are multiple components with anisotropic coefficients of thermal expansion, the internal strains and stresses are even more complex. Internal strains and stresses can arise in other ways as well.
Whatever their origin, the internal strains and stresses usually have adverse effects on the performance of the composite material. They often cause deterioration of the bonds between the components and laminates. The deterioration is manifested in lower measured property values than would be otherwise expected, and/or by observed bond line failure mechanisms. There may also be a shifting of the ultimate stress between the two dissimilar materials to a lower strength, more remote surface.
In a specific case of interest to the inventors, an aircraft structural member is fabricated by preparing a skin structure sub-assembly made of a precured quartz fiber/cyanate ester resin laminate and bonding the skin structure sub-assembly to a substrate sub-assembly formed of a low density (less than 25 pounds per cubic foot), surface-sealed syntactic foam, using an epoxy structural adhesive. This composite structure is observed to preferentially fail at the quartz fiber/cyanate ester resin interface. In many instances, failure occurred as the structure was cooled from the fabrication temperature to room temperature. Other bonded assemblies survived for a time at room temperature, but later failed catastrophically at the quartz fiber/cyanate ester resin interface. This failure mechanism indicates that the full strength potential of the skin structure and the syntactic foam is not realized because of the high stress developed as a result of the differences in the coefficients of thermal expansion of the two bonded sub-assemblies.
There is a need for an improved fabrication technique for such laminated composite structures. The present invention fulfills this need, and further provides related advantages.